1. Field of the Invention
The present invention relates to a method for filtering the vibratory excitations transmitted between two parts and an elastic linking device between two parts which transmits the static forces from one to the other in the axis of the device and for simultaneously filtering the associated coaxial vibratory excitations which are transmitted from one to the other.
More particularly, although not exclusively, such a device can be used in the suspension linking the main transmission gearbox to the fuselage of an aircraft with rotating wings, such as a helicopter, in order to filter the vibrations generated by the rotor and transmitted to the fuselage of said aircraft by said transmission gearbox.
2. Background Art
In fact, one of the fundamental problems of the helicopter arises from the general vibratory level which conditions, on the one hand, the level of the alternating stresses throughout the machine (and consequently the fatigue behavior and hence the lifetime of the parts) and, on the other hand, the cabin comfort and control vibrations.
The object of much research has therefore been to attenuate, if not to completely cancel out, this vibratory level inherent in the very operation of the rotor which not only transmits to the hub static or quasistatic forces and moments created by:
the lift (perpendicular to the plane of the rotor), PA1 the drag (in the plane of the rotor and parallel to the component V.sub.H, normal to the rotor mast, of the forward speed of the aircraft), PA1 the drift force (perpendicular to the preceding two and also in the plane of the rotor) which remains small and which can generally be ignored, but also periodic forces and moments originating from aerodynamic dissymetries (lift and profile drag) which appear during the rotation of the blades, essentially due to the forward speed in flight of translation at high speed, or also dissymetries resulting from the inequality of distribution of the speeds induced on the disk of the rotor at low speed (transition area). These alternate aerodynamic forces and moments are transmitted to the center of the rotor after having been attenuated or amplified by the blades. PA1 the forces (due to the flapping movements of the blades) and the moments (due to the drag movements of the blades) whose axes are carried by the axis of the rotor, are transmitted to the mast and to the fuselage only if their frequency expressed in hertz (Hz) is a harmonic of b.OMEGA., and thus of the form kb.OMEGA. (k : positive integer, equal to or greater than 1. The transfer of these forces and.. moments from the rotating axes to the fixed axes takes place without frequency change (oscillation and torsion effect in the structure); PA1 the forces (due to the drag movements of the blades) and the moments (due to the flapping movements of the blades) whose axes are in the plane of the rotor, are transmitted to the mast and to the fuselage only if their frequency is of the form (kb.+-.1).OMEGA., the resulting forces and moments then being at the frequency kb.OMEGA. in fixed axes (roll and pitch effects, transverse or longitudinal sway, principally in b.OMEGA.). PA1 the excitation harmonics at the site of the blades which affect the vibrations in the fuselage are distributed according to the order below: PA1 the higher the order of the harmonics, the smaller their amplitude; PA1 the excitations which affect comfort are the harmonics (kb.+-.1).OMEGA. in the axes of the blades found, after composition, at frequencies kb.OMEGA. in the fuselage of the helicopter; PA1 the lower the excitation frequency, the greater the extent to which people are sensitive to it, especially along the vertical axis. PA1 1) at the level of the rotor, with adoption of passive pendular antivibrators (or mass-spring systems arranged on the head of the rotor, for example) or of an active multi-cyclical control. In this latter case, a computer transmits signals to the cyclic pitch control of the blades by means of servocontrols, analyzes the effect produced and optimizes the input so as to have minimal acceleration as output. In other words, the multi-cyclical control is a solution specific to helicopters and based on the modification of the aerodynamic forces applied to the blades (and, consequently, on the modification of the excitation torque vector at the head of the rotor, thus at the exterior of the fuselage) by multicyclical injection of the pitch commands; PA1 2) at the level of the fuselage (designed at the outset with, among other imperatives, that of not having a characteristic vibration mode too close to the main excitation frequencies) by: PA1 a physical quantity which is representative of the vibratory excitations transmitted from one part to the other is measured, and corresponding first signals are generated, and PA1 said first signals are processed in order to convert them into second control signals for said actuator, which is controlled in order to oppose said vibratory excitations, PA1 said physical quantity is moreover measured on one or more of the additional passive links; and PA1 the dependent control signals for each of said actuators are derived by applying an automatic and continuous minimization of the performance criterion PI of formula: ##EQU1## in which: N=number of actuators and corresponding measurements, PA1 M=number of measurements on the additional passive link or links, PA1 .epsilon..sub.kf =harmonic component of rank k of said physical quantity of fundamental frequency f, PA1 p=number of harmonic components to be filtered, PA1 .vertline.a.vertline.=weighting matrix for the effect of each linking element. PA1 at least one linking element between said parts which comprises a transmission member for the static force between said parts and an actuator associated with said transmission member; PA1 at least one means of measurement of a physical quantity which is representative of the vibratory excitations transmitted from one part to the other, and able to supply corresponding first signals; and PA1 electronic processing means for said first signals in order to convert them into second control signals for said actuator, which is controlled in order to oppose said vibratory excitations. PA1 M=number of measurements on the additional passive link or links, PA1 .epsilon..sub.kf =harmonic component or rank k of said physical quantity of fundamental frequency f, PA1 p=number of harmonic components to be filtered, PA1 .vertline.a.vertline.=weighting matrix for the effect of each linking element.
Given that, in a general way, by .OMEGA. the speed of rotation of the rotor is expressed in number of revolutions per second and by b represents the number of blades, note that:
Consequently, note that a balanced rotor transmits, over and above the static forces and moments, only alternate forces and moments at a frequency which is a multiple of the speed of the rotor multiplied by the number of blades, the fundamental frequency being equal to b.OMEGA..
It would thus be appropriate, in order to avoid dangerous periodic forces at a frequency which is a multiple of the speed of the rotor, to increase the number of blades since:
______________________________________ two-blade 1 2 3 4 5 6 7 8 9 10 11 three-blade 2 3 4 5 6 7 8 9 10 11 four-blade 3 4 5 7 8 9 11 five-blade 4 5 6 9 10 11 ______________________________________
For reasons cost and of complexity, it is nevertheless appropriate to limit the number of blades.
Moreover, it is well known that as performance rises, the excitations increase as V.sup.n (n&gt;1): at high speed, the vibratory level in the fuselage grows in the same way.
These comments, together with evidence for ever more important comfort imperatives, are the justification for devising systems capable of transmitting the static forces and moments designated by F.sub.0 while attenuating the vibrations, which correspond to the decomposition into a Fourier series .SIGMA.Fn cos n.OMEGA.t (t: time, n: order of the harmonics). This attenuation must especially tend to minimize the vertical components of the dynamic loading at the level of the fuselage, which turn out to be the most troublesome in practice.
The parameters which condition the vibratory levels are accounted for of at the design stage of a rotor so as to minimize the effects thereof: type of hub (rigid or articulated hub) and choice of the number of blades, aerodynamic optimization of the blades in order to reduce the excitations and optimization of the dynamic response of the blades in order to reduce the torque vector (forces and moments) transmitted to the head of the rotor.
When these choices and the compromises that are reached do not yield the theoretical (or experimental) results anticipated for the vibratory levels, complementary means of action are resorted to in order to modify the excitation torque vector applied to the fuselage:
local processing, by modifying the shapes of the dynamic responses of the structure (batteries moved, stiffening of elements) or by installing mass-spring resonators (sprung batteries for example), PA2 overall processing, by active control of the structure, based on the modification of the excitation torque vector applied to the fuselage (that is to say the distribution of the interior forces) and of the response of the latter (French Patent Nos. 1,506,385 and 2,566,862); or PA2 3)intervention at the site of the interface between the mechanical assemblies of the main rotor and the fuselage in order to filter the transfer of the vibrations of the rotor to the airframe, especially via passive suspensions such as those described, for example, French Patent Nos. 1,507,306, 2,499,505 and 2,629,545.
More precisely, patent French Patent No. 1,506,385 relates to an attenuation method and an electrohydraulic attenuator for an aircraft with rotating wings. The method consists of creating, on the basis of the dynamic accelerations measured on the fuselage, electrical signals converted into variations in hydraulic pressure by means of an electrohydraulic servo valve, which pressure variations are transmitted to a double-acting jack arranged between the fuselage and the main transmission gearbox, in such a way as to oppose the vibrations. In order to do this, an accelerometer situated in the fuselage of the aircraft is linked to the control circuit of the double-acting jack. The jack constitutes a fourth linking bar, or, as an alternative, one of the usual bars which comprises, in this case, an elastic member in parallel which provides the flexibility necessary for the correct operation of the device while being able to take up the (substantially static) lift and maneuver forces.
A development of this concept is described in French Patent No. 2,566,862, in which, between the fuselage and the rotor of a helicopter, a plurality of actuators are provided whose oscillations are controlled in phase and in amplitude by virtue of the processing of signals which are representative of the dynamic accelerations, measured by a plurality of accelerometers arranged on the fuselage. Such a system operates in a closed loop. On the basis of the accelerometric measurements on the fuselage, the optimum commands to be generated are obtained with the use of a computer. The effective application of these commands modifies the condition of the fuselage and thus the subsequent measurements.
However, such accelerometric measurements are likely to be affected by errors and uncertainties, due especially to possible phase offsets, by the very fact that they are carried out on the fuselage of the aircraft.